Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method

ABSTRACT

A turbine engine includes a plurality of variable fan inlet guide vanes. Where the turbine engine is a tip turbine engine, the variable fan inlet guide vanes permit the ability to control engine stability even though the fan-turbine rotor assembly is directly coupled to the axial compressor at a fixed rate. The fan inlet guide vanes may be actuated from an inner diameter of the fan inlet guide vanes.

This application is a divisional of co-pending application U.S. Ser. No.13/022,456, filed Feb. 7, 2011, which is a divisional of U.S. Ser. No.11/719,143, filed May 11, 2007, now U.S. Pat. No. 7,882,694.

This invention was conceived in performance of U.S. Air Force contractF33657-03-C-2044. The government may have rights in this invention.

BACKGROUND OF THE INVENTION

The present invention relates to turbine engines, and more particularlyto a variable fan inlet guide vane for a turbine engine, such as a tipturbine engine.

An aircraft gas turbine engine of the conventional turbofan typegenerally includes a forward bypass fan, a low pressure compressor, amiddle core engine, and an aft low pressure turbine, all located along acommon longitudinal axis. A high pressure compressor and a high pressureturbine of the core engine are interconnected by a high spool shaft. Thehigh pressure compressor is rotatably driven to compress air enteringthe core engine to a relatively high pressure. This high pressure air isthen mixed with fuel in a combustor, where it is ignited to form a highenergy gas stream. The gas stream flows axially aft to rotatably drivethe high pressure turbine, which rotatably drives the high pressurecompressor via the high spool shaft. The gas stream leaving the highpressure turbine is expanded through the low pressure turbine, whichrotatably drives the bypass fan and low pressure compressor via a lowspool shaft.

Although highly efficient, conventional turbofan engines operate in anaxial flow relationship. The axial flow relationship results in arelatively complicated elongated engine structure of considerable lengthrelative to the engine diameter. This elongated shape may complicate orprevent packaging of the engine into particular applications.

A recent development in gas turbine engines is the tip turbine engine.Tip turbine engines may include a low pressure axial compressordirecting core airflow into hollow fan blades. The hollow fan bladesoperate as a centrifugal compressor when rotating. Compressed coreairflow from the hollow fan blades is mixed with fuel in an annularcombustor, where it is ignited to form a high energy gas stream whichdrives the turbine that is integrated onto the tips of the hollow bypassfan blades for rotation therewith as generally disclosed in U.S. PatentApplication Publication Nos.: 20030192303; 20030192304; and 20040025490.The tip turbine engine provides a thrust-to-weight ratio equivalent toor greater than conventional turbofan engines of the same class, butwithin a package of significantly shorter length.

Some low bypass ratio conventional turbine engines include variable faninlet guide vanes. The variable fan inlet guide vanes each include apivotably mounted flap. The trailing edges of the flaps are allconnected via activation levers to a unison ring about the outercircumference of the flaps, such that rotation of the unison ring causesthe flaps to pivot uniformly. Generally, high bypass ratio turbineengines (i.e. with a bypass ratio greater than three) do not includevariable fan inlet guide vanes.

SUMMARY OF THE INVENTION

A turbine engine according to the present invention includes a pluralityof variable fan inlet guide vanes. The variable fan inlet guide vanespermit the ability to control engine stability even though thefan-turbine rotor assembly is directly coupled to the axial compressorat a fixed rate. The variable fan inlet guide vane also provides lowerstarting power requirements and improved fan stability.

According to one feature disclosed herein, the fan inlet guide vanes areactuated from an inner diameter of the fan inlet guide vanes. Becausethe outer diameter of the fan inlet guide vane flaps would be verylarge, an outer diameter unison ring would deflect between the point atwhich the actuator contacts the unison ring and the diametricallyopposite flap, thereby reducing the uniformity of the flap actuation. Anouter diameter unison ring would also be very heavy.

BRIEF DESCRIPTION OF THE DRAWINGS

Other advantages of the present invention can be understood by referenceto the following detailed description when considered in connection withthe accompanying drawings wherein:

FIG. 1 is a partial sectional perspective view of a tip turbine engine.

FIG. 2 is a longitudinal sectional view of the tip turbine engine ofFIG. 1 along an engine centerline and a schematic view of an enginecontroller.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 illustrates a general perspective partial sectional view of a tipturbine engine (TTE) type gas turbine engine 10. The engine 10 includesan outer nacelle 12, a rotationally fixed static outer support structure14 and a rotationally fixed static inner support structure 16. Aplurality of fan inlet guide vanes 18 are mounted between the staticouter support structure 14 and the static inner support structure 16.Each fan inlet guide vane preferably includes a pivotable flap 18A. Anosecone 20 is preferably located along the engine centerline A toimprove airflow into an axial compressor 22, which is mounted about theengine centerline A behind the nosecone 20.

A fan-turbine rotor assembly 24 is mounted for rotation about the enginecenterline A aft of the axial compressor 22. The fan-turbine rotorassembly 24 includes a plurality of hollow fan blades 28 to provideinternal, centrifugal compression of the compressed airflow from theaxial compressor 22 for distribution to an annular combustor 30 locatedwithin the rotationally fixed static outer support structure 14.

A turbine 32 includes a plurality of tip turbine blades 34 (two stagesshown) which rotatably drive the hollow fan blades 28 relative aplurality of tip turbine stators 36 which extend radially inwardly fromthe rotationally fixed static outer support structure 14. The annularcombustor 30 is disposed axially forward of the turbine 32 andcommunicates with the turbine 32.

Referring to FIG. 2, the rotationally fixed static inner supportstructure 16 includes a splitter 40, a static inner support housing 42and a static outer support housing 44 located coaxial to said enginecenterline A.

The axial compressor 22 includes the axial compressor rotor 46, which ismounted for rotation upon the static inner support housing 42 through anaft bearing assembly 47 and a forward bearing assembly 48. A pluralityof compressor blades 52 extends radially outwardly from the axialcompressor rotor 46. A fixed compressor case 50 is mounted within thesplitter 40. The axial compressor 22 includes a plurality of inlet guidevanes 51 (one shown). For reasons explained below, it is not necessaryto provide a variable inlet geometry to the axial compressor 22.Therefore, the inlet guide vane 51 is fixed, thereby reducing the weightand complexity of the axial compressor 22.

A plurality of compressor vanes 54 extends radially inwardly from thecompressor case 50 between stages of the compressor blades 52. Thecompressor blades 52 and compressor vanes 54 are arrangedcircumferentially about the axial compressor rotor 46 in stages (threestages of compressor blades 52 and compressor vanes 54 are shown in thisexample).

The rotational position of the fan inlet guide vane flap 18A iscontrolled by an actuator 55 that is mounted within the nacelle 12,radially outwardly of one of the fan inlet guide vanes 18 and radiallyoutward of the bypass airflow path. The actuator 55 may be hydraulic,electric motor or linear actuator, or any other type of suitableactuator. The actuator 55 is operatively connected to the fan inletguide vane flaps 18A via a torque rod 56 that is routed through one ofthe inlet guide vanes 18. Within the splitter 40, the torque rod 56 iscoupled to a unison ring 57 via a torque rod lever 58. The unison ring57 is rotatable about the engine centerline A. The unison ring 57 iscoupled to a shaft 63 of the variable guide vane flap 18 a via anactivation lever 59. The plurality of variable guide vanes 18 and flaps18 a (only one shown) are disposed circumferentially about the enginecenterline A, and each is connected to the unison ring 57 in the samemanner. The actuator 55 is coupled to the torque rod 56 by an actuatorlever 60.

The fan-turbine rotor assembly 24 includes a fan hub 64 that supports aplurality of the hollow fan blades 28. Each fan blade 28 includes aninducer section 66, a hollow fan blade section 72 and a diffuser section74. The inducer section 66 receives airflow from the axial compressor 22generally parallel to the engine centerline A and turns the airflow froman axial airflow direction toward a radial airflow direction. Theairflow is radially communicated through a core airflow passage 80within the fan blade section 72 where the airflow is centrifugallycompressed. From the core airflow passage 80, the airflow is diffusedand turned once again by the diffuser section 74 toward an axial airflowdirection toward the annular combustor 30. Preferably, the airflow isdiffused axially forward in the engine 10, however, the airflow mayalternatively be communicated in another direction.

The tip turbine engine 10 may optionally include a gearbox assembly 90aft of the fan-turbine rotor assembly 24, such that the fan-turbinerotor assembly 24 rotatably drives the axial compressor 22 via thegearbox assembly 90. In the embodiment shown, the gearbox assembly 90provides a speed increase at a 3.34-to-one ratio. The gearbox assembly90 is an epicyclic gearbox, such as a planetary gearbox as shown, thatis mounted for rotation between the static inner support housing 42 andthe static outer support housing 44. The gearbox assembly 90 includes asun gear 92, which rotates the axial compressor rotor 46, and a planetcarrier 94, which rotates with the fan-turbine rotor assembly 24. Aplurality of planet gears 93 each engage the sun gear 92 and arotationally fixed ring gear 95. The planet gears 93 are mounted to theplanet carrier 94. The gearbox assembly 90 is mounted for rotationbetween the sun gear 92 and the static outer support housing 44 througha gearbox forward bearing 96 and a gearbox rear bearing 98. The gearboxassembly 90 may alternatively, or additionally, reverse the direction ofrotation and/or may provide a decrease in rotation speed.

A plurality of exit guide vanes 108 are located between the static outersupport housing 44 and the rotationally fixed exhaust case 106 to guidethe combined airflow out of the engine 10. An exhaust mixer 110 mixesthe airflow from the turbine blades 34 with the bypass airflow throughthe fan blades 28.

An upstream pressure sensor 130 measures pressure upstream of the fanblades 28 and a downstream pressure sensor 132 measures pressuredownstream of the fan blades 28. A rotation speed sensor 134 is mountedadjacent the fan blades 28 to determine the rotation speed of the fanblades 28. The rotation speed sensor 134 may be a proximity sensordetecting the passage of each fan blade 28 to calculate the rate ofrotation.

Control of the tip turbine engine 10 is provided by a Full AuthorityDigital Engine Controller (FADEC) 112 and by a fuel controller 114, bothmounted remotely from the tip turbine engine 10 (i.e. outside thenacelle 12) and connected to the tip turbine engine 10 by a singlewiring harness 116 and a single fuel line 118, respectively. The FADEC112 includes a power source 120 such as a battery, a fuel cell, or otherelectric generator. The FADEC 112 includes a CPU 122 and memory 124 forexecuting control algorithms to generate control signals to the tipturbine engine 10 and the fuel controller 114 based upon input from theupstream pressure sensor 130, the downstream pressure sensor 132 and therotation speed sensor 134. The control signals may include signals forcontrolling the position of the flaps 18A of the fan inlet guide vanes18, commands that are sent to the fuel controller 114 to indicate theamount of fuel that should be supplied and other necessary signals forcontrolling the tip turbine engine 10.

The fuel controller 114 also includes a power source 138, such as abattery, fuel cell, or other electric generator. The fuel controller 114includes at least one fuel pump 140 for controlling the supply of fuelto the tip turbine engine 10 via fuel line 118.

During operation, core airflow enters the axial compressor 22, where itis compressed by the compressor blades 52. The compressed air from theaxial compressor 22 enters the inducer section 66 in a directiongenerally parallel to the engine centerline A, and is then turned by theinducer section 66 radially outwardly through the core airflow passage80 of the hollow fan blades 28. The airflow is further compressedcentrifugally in the hollow fan blades 28 by rotation of the hollow fanblades 28. From the core airflow passage 80, the airflow is turned anddiffused axially forward in the engine 10 by the diffuser section 74into the annular combustor 30. The compressed core airflow from thehollow fan blades 28 is mixed with fuel in the annular combustor 30 andignited to form a high-energy gas stream.

The high-energy gas stream is expanded over the plurality of tip turbineblades 34 mounted about the outer periphery of the fan-turbine rotorassembly 24 to drive the fan-turbine rotor assembly 24, which in turnrotatably drives the axial compressor 22 either directly or via theoptional gearbox assembly 90. The fan-turbine rotor assembly 24discharges fan bypass air axially aft to merge with the core airflowfrom the turbine 32 in the exhaust case 106.

The FADEC 112 controls bypass air flow and impingement angle by varyingthe fan inlet guide vane flaps 18A based upon information in signalsfrom the upstream pressure sensor 130, the downstream pressure sensor132 and the rotation speed sensor 134. The sensors 130, 132, 134indicate a current operating state of the tip turbine engine 10. TheFADEC 112 determines a desired operating state for the tip turbineengine 10 and generates control signals to bring the tip turbine engine10 toward the desired operating state. These control signals includecontrol signals for varying the fan inlet guide vanes 18.

Closing the fan inlet guide vane flaps 18A during starting of the tipturbine engine 10 reduces the starter power requirements, whilemaintaining core airflow. During operation, the FADEC 112 controls theaxial compressor 22 operability and stability margin by varying the faninlet guide vane flaps 18A. In the tip turbine engine 10, the fan blades28 are coupled to the axial compressor 22 at a fixed rate via thegearbox 90 (or, alternatively, directly). Therefore, slowing therotation of the fan blades 28 by closing the fan inlet guide vane flaps18A slows rotation of the axial compressor 22. Additionally,controllably slowing down rotation of the fan blades 28 also reduces thecentrifugal compression of the core airflow in the fan blades 28 headingtoward the combustor 30, which thereby reduces the output of thecombustor 30 and the force with which the turbine 32 is rotated. Bysignificantly altering the speed-flow relationship of the primarypropulsor, the combustor temperature relationship changes in a way thatallows control of the primary compressor operating lines. This is drivenby the relationship between compressor exit corrected flow andhigh-pressure turbine inlet corrected flow. In typical gas turbineengines, the high-pressure turbine is typically choked and operates at aconstant inlet corrected flow. This combined with the fact that flow isusually proportional to speed and combustor temperature ratio istypically constant drives primary compressors to require some sort ofvariable geometry or bleed to maintain stability. By altering the fanspeed-flow characteristic through use of the fan variable inlet guidevanes 18, one can significantly alter the combustor temperature ratio,thereby controlling the primary compressor operating lines andestablishing stability without compressor variable geometry or bleed.

FIGS. 1 and 2 are generally scale drawings. The tip turbine engine 10shown is a high-bypass ratio turbine engine, with a bypass ratio of 5.0.

In accordance with the provisions of the patent statutes andjurisprudence, exemplary configurations described above are consideredto represent a preferred embodiment of the invention. However, it shouldbe noted that the invention can be practiced otherwise than asspecifically illustrated and described without departing from its spiritor scope. Alphanumeric identifiers on method steps are for ease ofreference in dependent claims and do not signify a required sequence ofperformance unless otherwise indicated.

1. A method for controlling a tip turbine engine including the steps of:determining a desired change in operation of the tip turbine engine; andvarying a fan inlet geometry to the tip turbine engine in order toimplement the desired change in operation of the tip turbine engine. 2.The method of claim 1 wherein the tip turbine engine includes a fixedgeometry inlet to an axial compressor.
 3. The method of claim 1 whereinsaid step b) further includes the step of imparting a force on an innerradial end of a flap of a fan inlet guide vane.
 4. The method of claim 3wherein the fan inlet guide vane is one of a plurality of fan inletguide vanes.
 5. The method of claim 4 wherein said step b) furtherincluding the step of imparting the force on inner radial ends of flapson each of the fan inlet guide vanes.
 6. The method of claim 1 furtherincluding the step of centrifugally compressing core airflow within acentrifugal compression chamber inside a fan blade downstream of the faninlet geometry.
 7. The method of claim 1 wherein the tip turbine enginehas a high bypass ratio.
 8. The method of claim 1 wherein the tipturbine engine has a bypass ratio greater than three.
 9. The method ofclaim 8 further including the step of closing the fan inlet geometryduring starting of the tip turbine engine.
 10. The method of claim 1further including the step of slowing the rotation of a fan of the tipturbine engine in order to reduce compression of air into a combustor ofthe tip turbine engine during starting of the tip turbine engine.
 11. Amethod for controlling a high bypass ratio turbine engine including thesteps of: determining a desired change in operation of the turbineengine; and varying a fan inlet geometry to the turbine engine in orderto implement the desired change in operation of the turbine engine. 12.The method of claim 11 wherein said step b) further including the stepof imparting a force on inner radial ends of flaps on a plurality of faninlet guide vanes.
 13. The method of claim 12 wherein the turbine enginehas a bypass ratio greater than three.
 14. The method of claim 13further including the step of closing the fan inlet geometry duringstarting of the turbine engine.
 15. The method of claim 14 furtherincluding the step of slowing the rotation of a fan of the turbineengine in order to reduce compression of air into a combustor of theturbine engine during starting of the turbine engine.
 16. The method ofclaim 15 wherein the turbine engine does not include variable geometryinlet for an axial compressor.
 17. The method of claim 13 wherein theturbine engine does not include variable geometry inlet for an axialcompressor.